The present invention generally relates to a gas turbine engine, such as one used in aerospace or industrial application, in which a combustor with discrete fuel nozzles burns fuel and discharges hot gases into a conventional turbine engine. More particularly, the present invention relates to a method and apparatus for circumferential alignment of fuel nozzles and downstream turbine vanes on a gas turbine engine to significantly improve the hot-end durability for the static components.
A turbine engine normally includes stages of static airfoils, commonly called vanes, inter-spaced between stages of rotating airfoils, commonly called blades.
The combustor typically includes a discrete number of fuel nozzles, which deliver fuel and facilitate mixing with the air to achieve a stable, self-sustainable flame in the combustor chamber. Although it would be desired to have a uniform circumferential temperature field at the combustor exit, hotter and colder zones exist due to the discrete number of fuel nozzles. The hotter zones of combustion gas are commonly known as hot streaks. These hot streaks are known to pass through a first stage of vanes, the following blades, and on to the second stage of vanes. Similar interaction occurs between subsequent stages of the turbine engine.
More specifically, a hot-streak is a high temperature gas stream that passes through airfoil stages, both vanes and blades. The impact of having a hot-streak impinge on the stationary vane is a significant increase in airfoil metal temperature, leading to accelerated oxidation, thermo-mechanical fatigue, and a commensurate reduction in durability. This reduction in durability of the first-stage vane leads to frequent replacement at a high cost to turbine engine operators. When an airfoil is exposed to a higher temperature environment, the mechanical damage in the airfoil metal accelerates faster, causing the airfoil to lose its durability. Therefore, the durability of stationary vanes highly depends on the maximum hot-streak temperature. When a fuel nozzle is fouled by carbon, or if significant combustor flow field variability exists, the hot-streak temperature tends to increase and results in significant damage to the stationary vanes. Thus, vane durability can be improved if the vane airfoils are placed in generally cooler regions of combustion gas.
U.S. Pat. No. 5,486,091, Jan. 23, 1996 for Gas Turbine Airfoil Clocking, clocks the wakes from upstream stages of turbine airfoils to impinge the wakes on the leading edge of downstream stages of turbine airfoils. Clocking of the wakes specifically takes place at the long term operating condition. The ""091 claims that, at each successive stage of nozzle airfoils, the efficiency of the gas turbine engine is improved.
While a wake is an aerodynamic disturbance in the flow field, a hot streak is a temperature difference phenomenon. A wake is a zone of aerodynamic disturbance produced by a body such as an airfoil vane or blade placed in a flow stream. The wake is a low momentum region of fluid flow downstream of a body caused by the body""s profile and surface roughness. The body produces profile drag because the body exerts a force on the fluid, thus reducing the fluid momentum. The body also produces viscous drag due to the airfoil surface boundary layer, which also reduces the fluid momentum. The result of the profile and viscous drag is a region of low momentum fluid (the wake) which propagates downstream and causes pressure losses as the wake mixes with the free-stream fluid. Wake propagation to downstream stages is a result of mixing from shear forces due to velocity gradients in the flow field.
By contrast, a hot streak is a zone of hot fluid relative to the surrounding flow field. It requires no body (such as a vane or blade airfoil), is not a momentum deficit region like a wake, and is not associated with profile or viscous drag phenomenon. A hot streak is a temperature difference phenomenon, which propagates downstream by conduction and convection mixing of warmer and cooler fluids. Thus, tracing wake or hot streak propagation to downstream airfoil stages requires different analyses. To trace the hot streak, a non-uniform temperature profile must be imposed in the analysis model. The hot streak can then be traced by inviscid tools. Tracing a wake requires no specific inlet temperature profile. However, a viscous analysis or an inviscid analysis utilizing an artificial method to generate the wake must be used. A standard unsteady inviscid analysis will not be capable of tracing a wake but can be used to trace a hot streak. Both vector diagram and Computational Fluid Dynamics (CFD) analyses have shown that the fluid in a hot streak is accelerated through the vane and rotating blade stages differently than the surrounding cooler fluid, because it is less dense than the surrounding fluid (due to its increased temperature). The hot streak fluid thus accelerates to a higher velocity than the surrounding cooler fluid, resulting in a different vector diagram for the hot streak fluid. Therefore, a wake and a hot streak do not follow the same paths as they propagate to downstream airfoil stages. As a result, aligning downstream components by wake analysis is fundamentally different than aligning components by hot streak analysis. Further, an approach for improving engine efficiency is not necessarily the same as one for improving the durability of the hot section vanes, which involves hot streak temperature effects.
Moreover, the hotter operating temperature conditions, and thus the more severe hot streak damage to airfoils, typically occur at take-off and high temperature climb conditions, where turbine inlet temperatures are at their highest, and where the majority of the oxidation and thermo-mechanical fatigue damage can occur. A turbine engine aircraft may spend only a small fraction of its time at take-off and high temperature climb conditions. Therefore, optimization for the long term operating condition for turbine engine efficiency may not likely address hot streak damage and the durability of airfoils.
In an effort to control the temperature of hot gases at the combustor exit plane to enhance vane durability, U.S. Pat. No. 4,733,538, Mar. 29, 1988 for Combustion Selective Temperature Dilution, first establishes a pre-selected, desired temperature gradient exiting a combustor to purposely suppress the gas temperature in the region of the vane airfoil. By virtue of a specific placement of dilution air apertures, such pre-selected or favorable temperature gradient or distribution of high and low temperature zones is achieved. The dilution air apertures are aligned downstream of the fuel nozzles and axially aligned with the turbine vanes, as well as aligned with gaps between the vanes. One set of apertures directs some dilution air to a region physically close to the vane airfoils to permit such air to suppress the gas temperature near the vane airfoils. Another set of apertures directs some dilution air in between the vanes. Thereby, the temperature of the hot gases near the vane airfoils is reduced below the average gas temperature, while the temperature of the hot gases flowing between the vane airfoils is in excess of the average gas temperature. Accordingly, the temperature profile exiting the combustor is altered to obtain the preselected temperature profile to enhance durability for the existing placement of the vanes, and the placement of the vanes is not altered.
However, the placement of such dilution air apertures downstream in the combustor may be disadvantageous. For example, the dilution air may be better used in the combustor to meet stringent emissions and smoke requirements or to enhance durability of combustor components.
As may be seen from the foregoing discussion, there is a need for a method and apparatus that provides improved airfoil durability from alignment of combustor hot streaks to downstream turbine vanes.
In one aspect of the present invention, a method is provided to significantly reduce the average and maximum temperatures to which the turbine vane airfoils in the hot-section of a gas-turbine engine are subjected. This method relates to the circumferential alignment of fuel nozzles and downstream turbine vane airfoils on a gas turbine engine and positions the hot-streak emerging from each fuel nozzle in between the like-numbered of turbine vane airfoils. By employing computational fluid dynamic models, aerodynamic particle tracing using vector diagrams, or experimentally determined alignment from engine testing, information required for the circumferential alignment can be obtained for the vane airfoils to be placed away from the combustor periodic hot streaks. The advantage of this invention is to significantly improve the hot-end durability for the static components.
In another aspect of the present invention, a turbine engine embodies hot streaks emerging from fuel nozzles and directed in between the nozzle airfoils. For maximum airfoil durability, the number of vanes at different turbine stages is the same as the number of fuel nozzles. The vane airfoils in each turbine stage are circumferentially placed to avoid the high temperature zones generated by the combustor periodic hot streaks.
These and other features, aspects and advantages of the present invention will become better understood with reference to the following drawings, description and claims.